Auxiliary power unit with combined cooling of generator

ABSTRACT

An auxiliary power unit for an aircraft, including an internal combustion engine having a liquid coolant system, a generator drivingly engaged to the internal combustion engine and having a liquid coolant system distinct from the liquid coolant system of the internal combustion engine, a first heat exchanger in fluid communication with the liquid coolant system of the internal combustion engine, a second heat exchanger in fluid communication with the liquid coolant system of the generator, an exhaust duct in fluid communication with air passages of the heat exchangers, and a fan received in the exhaust duct and rotatable by the internal combustion engine for driving a cooling air flow through the air passages. The liquid coolant system of the engine may be distinct from fuel and lubricating systems of the auxiliary power unit. A method of cooling a generator and an internal combustion engine is also discussed.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a divisional of U.S. application Ser. No. 15/227,483filed on Aug. 3, 2016, which claims priority from U.S. application No.62/202,275 filed Aug. 7, 2015, the entire contents of which areincorporated by reference herein.

TECHNICAL FIELD

The application relates generally to compound engine assemblies, moreparticularly to such assemblies used as auxiliary power units (APU).

BACKGROUND OF THE ART

Traditional gas turbine engine auxiliary power units including an enginecore with a combustor which are used to drive a generator typicallyrequire a cooling system for the generator. Such a cooling system mayinclude fans and/or ejectors can represent significant power lossesand/or create drag penalties in flight.

Moreover, such traditional gas turbine engine auxiliary power unitsusually have an exhaust with relatively high temperature, requiring theuse of high temperature materials in the exhaust duct walls, which mayrepresent a significant cost.

SUMMARY

In one aspect, there is provided an auxiliary power unit for anaircraft, the auxiliary power unit comprising: an internal combustionengine having a liquid coolant system distinct from any fuel andlubricating system of the auxiliary power unit; a generator drivinglyengaged to the internal combustion engine, the generator having a liquidcoolant system distinct from the liquid coolant system of the internalcombustion engine; a first heat exchanger having first coolant passagesin fluid communication with the liquid coolant system of the internalcombustion engine and first air passages in heat exchange relationshipwith the first coolant passages; a second heat exchanger having secondcoolant passages in fluid communication with the liquid coolant systemof the generator and second air passages in heat exchange relationshipwith the second coolant passages; an exhaust duct in fluid communicationwith the first and second air passages; and a fan received in theexhaust duct and rotatable by the internal combustion engine for drivinga cooling air flow through the first and second air passages.

In another aspect, there is provided an auxiliary power unit for anaircraft, the auxiliary power unit comprising: an internal combustionengine having a liquid coolant system; a compressor having an outlet influid communication with an inlet of the internal combustion engine; aturbine section having an inlet in fluid communication with an outlet ofthe internal combustion engine, the turbine section including at leastone turbine compounded with the internal combustion engine; a generatordrivable by the internal combustion engine and having a liquid coolantsystem distinct from the liquid coolant system of the internalcombustion engine; a first heat exchanger in fluid communication withthe liquid coolant system of the internal combustion engine; a secondheat exchanger in fluid communication with the liquid coolant system ofthe generator; and a fan rotatable by the internal combustion engine fordriving a cooling air flow through the first and second heat exchangers.

In a further aspect, there is provided a method of cooling a generatorand an internal combustion engine of an auxiliary power unit for anaircraft, the method comprising: circulating a first liquid coolantthrough the internal combustion engine; circulating a second liquidcoolant through the generator; and driving a cooling air flow in heatexchange relationship with the first and second liquid coolants using afan driven by the internal combustion engine.

DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures in which:

FIG. 1 is a schematic cross-sectional side view of an auxiliary powerunit in accordance with a particular embodiment;

FIG. 2 is a schematic cross-sectional plan view of the auxiliary powerunit of FIG. 1;

FIG. 3 is a schematic tridimensional view of the auxiliary power unit ofFIG. 1;

FIG. 4 is a schematic cross-sectional view of a rotary engine which maybe used in the auxiliary power unit of FIGS. 1-3;

FIG. 5 is a schematic tridimensional view of an auxiliary power unit inaccordance with another particular embodiment;

FIG. 6 is another schematic tridimensional view of the auxiliary powerunit of FIG. 5, taken from an opposite side;

FIG. 7 is a schematic cross-sectional view of part of the auxiliarypower unit of FIG. 5;

FIG. 8 is a schematic tridimensional view, partly in transparency, of anend of the auxiliary power unit of FIG. 5 received in a tail cone of anaircraft;

FIG. 9 is a schematic bottom view of an auxiliary power unit and tailcone in accordance with a particular embodiment, with part of the tailcone removed for clarity;

FIG. 10 is a schematic side view of the auxiliary power unit and tailcone of FIG. 9, with part of the tail cone removed for clarity;

FIG. 11 is a schematic cross-sectional view of compressor and turbinesections of the auxiliary power units of FIG. 5 and of FIG. 9;

FIG. 12 is a schematic cross-sectional view of part of an auxiliarypower unit showing a cooling inlet and heat exchanger configuration inaccordance with another particular embodiment which may be alternatelyused in any of the above auxiliary power units;

FIG. 13 is a schematic cross-sectional view of part of an auxiliarypower unit showing a cooling inlet and heat exchanger configuration inaccordance with another particular embodiment which may be alternatelyused in any of the above auxiliary power units;

FIG. 14 is a schematic cross-sectional view of a compressor section inaccordance with another particular embodiment which may be alternatelyused in any of the above auxiliary power units;

FIG. 15 is a diagram of compressor and turbine configuration inaccordance with another particular embodiment which may be alternatelyused in any of the above auxiliary power units; and

FIG. 16 is a schematic cross-sectional view of the compressor andturbine configuration of FIG. 15.

DETAILED DESCRIPTION

The present description includes compound engine assembly auxiliarypower units for providing supplementary ground and flight pneumaticand/or electric power for airborne auxiliary power unit applications. Ina particular embodiment, the auxiliary power units are configured todirectly replace a traditional gas turbine engine auxiliary power unitand perform in a more efficient manner, with power/weight andpower/volume properties meeting the requirements for airborneapplication. Application to fixed or mobile ground power units is alsopossible.

Referring to FIGS. 1-3, an auxiliary power unit 10 in accordance with aparticular embodiment is generally shown. The auxiliary power unit 10includes an engine core 12′ including one or more intermittent internalcombustion engines 12 engaged to a common shaft 16 (see FIG. 2). In aparticular embodiment, the intermittent internal combustion engine(s) 12is/are rotary internal combustion engine(s), for example Wankelengine(s); it is however understood that other types of intermittentinternal combustion engines may alternately be used.

Referring to FIG. 4, an example of a Wankel engine which may be used inthe engine core 12′ is shown. It is understood that the configuration ofthe engine(s) 12, e.g. placement of ports, number and placement ofseals, etc., may vary from that of the embodiment shown. The engine 12comprises a housing 32 defining a rotor cavity having a profile definingtwo lobes, which is preferably an epitrochoid. A rotor 34 is receivedwithin the rotor cavity. The rotor defines threecircumferentially-spaced apex portions 36, and a generally triangularprofile with outwardly arched sides. The apex portions 36 are in sealingengagement with the inner surface of a peripheral wall 38 of the housing32 to form and separate three working chambers 40 of variable volumebetween the rotor 34 and the housing 32. The peripheral wall 38 extendsbetween two axially spaced apart end walls 54 to enclose the rotorcavity.

The rotor 34 is engaged to an eccentric portion 42 of an output shaft 16to perform orbital revolutions within the rotor cavity. The output shaft16 performs three rotations for each orbital revolution of the rotor 34.The geometrical axis 44 of the rotor 34 is offset from and parallel tothe axis 46 of the housing 32. During each orbital revolution, eachchamber 40 varies in volume and moves around the rotor cavity to undergothe four phases of intake, compression, expansion and exhaust.

An intake port 48 is provided through the peripheral wall 38 foradmitting compressed air into one of the working chambers 40. An exhaustport 50 is also provided through the peripheral wall 38 for discharge ofthe exhaust gases from the working chambers 40. Passages 52 for a sparkplug, glow plug or other ignition mechanism, as well as for one or morefuel injectors of a fuel injection system (not shown) are also providedthrough the peripheral wall 38. Alternately, the intake port 48, theexhaust port 50 and/or the passages 52 may be provided through the endor side wall 54 of the housing. A subchamber (not shown) may be providedin communication with the chambers 40, for pilot or pre injection offuel for combustion.

For efficient operation the working chambers 40 are sealed byspring-loaded peripheral or apex seals 56 extending from the rotor 34 toengage the inner surface of the peripheral wall 38, and spring-loadedface or gas seals 58 and end or corner seals 60 extending from the rotor34 to engage the inner surface of the end walls 54. The rotor 34 alsoincludes at least one spring-loaded oil seal ring 62 biased against theinner surface of the end wall 54 around the bearing for the rotor 34 onthe shaft eccentric portion 42.

The fuel injector(s) of the engine 12, which in a particular embodimentare common rail fuel injectors, communicate with a source of Heavy fuel(e.g. diesel, kerosene (jet fuel), equivalent biofuel), and deliver theheavy fuel into the engine 12 such that the combustion chamber isstratified with a rich fuel-air mixture near the ignition source and aleaner mixture elsewhere.

Referring back to FIGS. 1-3, the auxiliary power unit 10 includes asupercharger compressor 20 having an outlet in fluid communication withthe inlet of the engine core 12′ (e.g. intake port 48 of each engine12). Air enters an inlet plenum 19 from the aircraft inlet 14, and theair is compressed by the compressor 20 which optionally includesvariable inlet guide vanes 23 and optionally includes a variablediffuser 25 (FIG. 2), which in a particular embodiment allow formanagement of a wide range of flow and pressure ratio conditions. Theair from the compressor 20 circulates through an intercooler heatexchanger 18 to drop its temperature, for example from about 450° F. to250° F., prior to entering the engine core. In the embodiment shown, thecompressor 20 also provides bleed air for the aircraft; after leavingthe compressor 20 and before reaching the intercooler 18, a portion ofthe compressed air is directed to a bleed duct 27 to be delivered to theaircraft.

At certain operating conditions it may be necessary to bleed excess airfrom the compressor 20 to avoid surge. In the embodiment shown, theconduit between the compressor 20 and the intercooler 18 is in fluidcommunication with an excess air duct 29 to bleed this excess air; adiverter valve 31 is incorporated in the excess air duct 29 to managethe flow of air being bled from the compressor 20. The diverter valve 31may be scheduled to open based on sensed compressor exit conditionsindicating operation close to surge.

In the engine core 12′ air is mixed with fuel and combusted to providepower and a residual quantity of intermediate pressure exhaust gas. Theoutlet of the engine core 12′ (e.g. exhaust port 50 of each engine 12)is in fluid communication with an inlet of a turbine section, so thatthe exhaust gases from the engine core 12′ are expanded in the turbinesection. The turbine section has one or more turbines 26, 22 compoundedwith the engine core 12′. In a particular embodiment, the turbinesection includes a first stage turbine 26 having an outlet in fluidcommunication with an inlet of a second stage turbine 22, with theturbines 26, 22 having different reaction ratios from one another. Thedegree of reaction of a turbine can be determined using thetemperature-based reaction ratio (equation 1) or the pressure-basedreaction ratio (equation 2), which are typically close to one another invalue for a same turbine, and which characterize the turbine withrespect to “pure impulse” or “pure reaction” turbines:

$\begin{matrix}{{{Reaction}\; (T)} = \frac{( {t_{S\; 3} - t_{S\; 5}} )}{( {t_{S\; 0} - t_{S\; 5}} )}} & (1) \\{{{Reaction}\; (P)} = \frac{( {P_{S\; 3} - P_{S\; 5}} )}{( {P_{S\; 0} - P_{S\; 5}} )}} & (2)\end{matrix}$

where t is temperature and P is pressure, s refers to a static port, andthe numbers refers to the location the temperature or pressure ismeasured: 0 for the inlet of the turbine vane (stator), 3 for the inletof the turbine blade (rotor) and 5 for the exit of the turbine blade(rotor); and where a pure impulse turbine would have a ratio of 0 (0%)and a pure reaction turbine would have a ratio of 1 (100%).

In a particular embodiment, the first stage turbine 26 is configured totake benefit of the kinetic energy of the pulsating flow exiting thecore engine(s) 12 while stabilizing the flow and the second stageturbine 22 is configured to extract energy from the remaining pressurein the flow. Accordingly, in a particular embodiment the first stageturbine 26 has a lower reaction ratio (i.e. lower value) than that ofthe second stage turbine 22. In a particular embodiment, the first stageturbine 26 has a reaction ratio of 0.25 or lower (temperature orpressure based) or of 0.2 or lower (temperature or pressure based), andthe second stage turbine 22 a reaction ratio higher than 0.25(temperature or pressure based) and/or is a medium reaction pressureturbine. Other values are also possible.

The compressor 20 may be driven by one or more of the turbines 26, 22and/or the engine core 12; in the embodiment shown and as can be bestseen in FIG. 2, the first and second stage turbines 26, 22 andcompressor 20 are coupled to the same shaft 24. In a particularembodiment, the turbines 26, 22 and compressor 20 coupled on the sameshaft 24 allow for a reasonably efficient non dimensional specific speedmatch between the compressor and turbine section. In a particularembodiment the turbine shaft 24 rotates at approximately 40,000 to50,000 rpm; other values for the rotational speeds are also possible.

In the embodiment shown, the first and second stage turbines 26, 22 areboth compounded with the engine core 12′ by having the turbine andengine shafts 24, 16 coupled through a gearbox 28. In a particularembodiment, the transmission of the gearbox 28 includes a compound geartrain such that torque and power may be communicated between the turbineand engine shafts 24, 16 in either direction.

In a particular embodiment, part of the compressor airflow which isdelivered to the aircraft forms the output “load”. A large part of thisload is supported by the turbines 26, 22 on the same shaft 24 andtherefore the load on the engine core 12′ transmitted via the gearbox 28is minimized. Thus losses and additional heat from the gearbox 28 may beminimized. Alternatively if the turbines 26, 22 provide more power thanthe compressor 20 requires the excess torque transmitted to the enginecore 12′ may be relatively small.

In a particular embodiment, the engine core 12′ including rotaryinternal combustion engine(s) 12 runs at approximately 8000 rpm; othervalues are also possible. In a particular embodiment, the combined stepup gear ratio defined by the gearbox 28 between the engine core shaft 16and the turbine shaft 24 is between about 4:1 and 7:1, for example about5:1. In a particular embodiment, a two stage compound idler system isused to provide the appropriate ratio and provide offset centres betweenthe engine core shaft 16 and the turbine shaft 24. The offset betweenthe engine core shaft 16 and the turbine shaft 24 may allow for the hotexhaust output from the ports 50 of the core engines 12 to be ducteddirectly into the turbine section while minimizing the length of theducts.

A generator 64 is drivable by the engine core 12′ to provide aircraftelectrical power for accessories and/or control purposes, for example bybeing driven through mechanical engagement with the engine core 12′directly or through the gearbox 28, or by mechanical engagement with theturbine shaft 24. In the embodiment shown, the generator 64 is mounteddirectly (i.e. without intermediate gearing) to the end of the enginecore shaft 16. In a particular embodiment the generator 64 is a 400 Hz,6 pole alternator/generator with a design synchronous speed of 8000 rpm;other configurations are also possible. The alternator/generator 64 mayserve as a startor. In a particular embodiment, elimination of anyintermediate gearing between the engine core shaft 16 and thealternator/generator 64 eliminates heat generation and loss associatedwith that gearing (which may generally corresponds to approximately 2%of the rated generator load).

In a particular embodiment, the auxiliary power unit 10 includes a fullauthority electronic control managing all the operational requirements.The control system manages the compressor inlet guide vanes 23 and/orvariable diffuser 25 (if applicable) of the shared supercharger andaircraft bleed compressor 20 to achieve the required bleed pressure andflow to the bleed duct 27 and the required fuel/air ratio in the enginecore 12′ to maintain the governed speed. In the event of conflictbetween the aircraft air requirements and the governed speed, thecompressor variables are set as required to allow the system to maintainthe governed speed and provide priority to the generator power. In theevent this action causes excess air flow or excess pressure, theseconditions may be managed by opening the diverter valve 31. A load valve(not shown) can also optionally be provided in the bleed duct 27 andmanaged by the control system to throttle or cut off the air supply tothe aircraft.

With a constant volume combustion cycle in the engine core 12′ thebreakdown of waste heat of the auxiliary power unit 10 is different froma traditional gas turbine engine auxiliary power unit. Less heat isevacuated through the exhaust and more heat is given up to the enginecasing. Accordingly, the engine(s) 12 of the engine core 12′ have acoolant system which in a particular embodiment is distinct from anyfuel and lubricating system of the auxiliary power unit 10; in otherwords, a dedicated coolant is circulated through the engine(s) 12 of theengine core 12′, for example through multiple coolant passages definedin the walls of the housing 32, and this dedicated coolant is circulatedseparately and independently from the lubricant and the fuel of theauxiliary power unit 10, including the lubricant of the engine core 12′.The dedicated coolant may be a liquid coolant, for example water. A heatexchanger defining an engine core cooler 66 includes coolant passages 66a (see FIG. 1) in fluid communication with the coolant system of theengine core 12′ and air passages 66 b (see FIG. 1) in heat exchangerelationship with the coolant passages 66 a.

The generator 64 also includes a coolant system distinct from thecoolant system of the engine(s) 12; the coolant system of the generatormay be independent from or may be common with a lubrication system ofthe generator 64. The generator coolant may be a liquid coolant, forexample oil. A second heat exchanger defining a generator cooler 68includes coolant passages 68 a (see FIG. 1) in fluid communication withthe coolant system of the generator 64 and air passages 68 b (seeFIG. 1) in heat exchange relationship with the coolant passages 68 a. Inthe embodiment shown, both coolers 66, 68 are provided in a commonpackage, with the coolant passages 66 a, 68 a of the two coolers 66, 68being distinct from one another. In a particular embodiment where thegenerator coolant is oil or another suitable lubricant, the generatorcoolant system is common with (in fluid communication with) thelubrication system of the auxiliary power unit 10, which distributeslubricant to various components of the auxiliary power unit 10 (e.g.bearings, gears, etc., of the engine core 12′, the compressor 20, theturbines 22, 26, the gearbox 28), so the second heat exchanger 68 isalso an engine lubricant cooler. Alternately, a separate heat exchanger(not shown) may be provided for the lubrication system of the auxiliarypower unit 10, and the cooler 68 may be configured to cool only thegenerator lubricant/coolant.

The air passages 66 b, 68 b of the coolers 66, 68 are in fluidcommunication with an exhaust duct 70 of the auxiliary power unit 10;the exhaust duct 70 has an outlet 72 in fluid communication with theenvironment of the aircraft, so that the cooling air flow can bedischarged to atmosphere. The exhaust duct 70 defines a cooling inlet 74in fluid communication with an aircraft compartment 76 containing theauxiliary power unit 10. In the embodiment shown, the coolers 66, 68 arereceived in the exhaust duct 70. The intercooler 18 is also received inthe exhaust duct 70, upstream of the coolers 66, 68.

A fan 78 (FIG. 2) is rotatable by the engine core 12′ and in fluidcommunication with the exhaust duct 70 for driving the cooling air flowfrom the compartment 76, through the heat exchangers (coolers 66, 68 andintercooler 18) and out of the exhaust duct 70 to atmosphere. In theembodiment shown, the fan 78 is received in the exhaust duct 70 upstreamof the heat exchangers 18, 66, 68 and is directly driven by the enginecore 12′, by being mounted on the end of the engine core shaft 16opposite from the generator 64. In a particular embodiment, direct driveof the fan 78 by the engine core shaft 16 allows to avoid additionalgear loss and heat which would be produced by a gear drive. Alternately,the fan 78 may be driven through a transmission (whether in gearbox 28or another transmission specific to the fan 78), or be electrically orhydraulically driven by a motor obtaining power directly or indirectlyfrom the engine core 12′.

In a particular embodiment, the blade speed of the fan 78 issufficiently low such that the fan 78 can be made of a common Al alloy,organic composite or thermoplastic material. In a particular embodiment,the fan 78 rotates at about 8000 rpm; other values are also possible.

Rotation of the fan 78 induces flow from the compartment 76, which alsoprovides a compartment ventilation function. In a particular embodiment,side openings from the main aircraft inlet 14 allow cooling air to flowinto the compartment 76 under the driving action of the fan 78 to coolthe surfaces of the auxiliary power unit 10 exposed within thecompartment 76. In a particular embodiment, the fan inlet is protectedby a screen to prevent larger objects from damaging the fan 78.

Although multiple distinct coolers are shown in series on FIGS. 1-3,alternately only one integrated cooler unit may be used with areas subdivided dedicated to the engine lubricant/generator coolant, engine coreliquid coolant, and inter-cooling functions. The heat exchangers 18, 66,68 may also be angled at an angle more than 90° to the flow direction,for example to optimize the area presented to the airflow. Although notshown, the coolers 66, 68 may include a thermal bypass system to preventover-cooling at lower ambient temperatures, for example managed by theelectronic control system based on sensed coolant temperatures, or byany other suitable thermostat concept.

The cooling system of the engine core 12′ is thus integrated with thatof the generator 64 and with the cooling system for the lubricant of theauxiliary power unit 10. In a particular embodiment, this integrationallows for a reduction or minimization of the power loss from fans andejectors traditionally used, and/or to avoid cooling drag penalties inflight. In a particular embodiment, the auxiliary power unit 10 isconfigured to reduce or avoid the generation of additional heat, forexample from gear train losses.

Through the integrated cooling system, the same fan 78 drives thecooling air flow through the compartment 76, engine core cooler 66,intercooler 18, and generator/engine lubricant cooler 68, and thendischarges the cooling air out to atmosphere through the exhaust duct70; in a particular embodiment, the entire auxiliary power unit 10 andits cooling system can be installed and removed as a single assemblywith interconnects and aircraft inlet and exhaust similar to that of atraditional gas turbine engine auxiliary power unit. In use and in aparticular embodiment, the generator 64 and the engine core 12′ are thuscooled by circulating a first coolant (e.g. water) through the engine(s)12 of the engine core 12′, circulating a second coolant (e.g. oil)through the generator 64, and driving the cooling air flow in heatexchange relationship with the first and second coolants using the fan78 driven by the auxiliary power unit 10.

If applicable any diverted air from the compressor 20 can also beintroduced in the exhaust duct 70. Accordingly, in the embodiment shown,the excess air duct 29 provides a direct fluid communication between thecompressor 20 and a portion of the exhaust duct 70 located downstream ofthe fan 78 and heat exchangers 18, 66, 68.

In a particular embodiment, the exhaust duct 70 is located in a tailcone of the aircraft. As can be best seen in FIGS. 1-2, an intermediateduct 80 extends in fluid communication with the exhaust of the enginecore 12′, by being connected to an exhaust of the second stage turbine22. The intermediate duct 80 has an outlet 82 positioned in the exhaustduct 70, downstream of the fan 78 and upstream of the outlet 72 of theexhaust duct 70. The outlet 82 of the intermediate duct 80 is spacedradially inwardly from a peripheral wall 70′ of the exhaust duct 70. Theair and exhaust gases are thus discharged in the exhaust duct 70 so thatthe flow of cooling air surrounds the flow of exhaust gases. The massflow and/or volume of flow of exhaust gases is/are smaller than the flowof cooling air. In a particular embodiment, the mass flow of exhaustgases is 20% or less of the mass flow of cooling air. An opencross-sectional area of the outlet 82 of the intermediate duct 80 issmaller than an open cross-sectional area of the exhaust duct 70 aroundthe outlet 82 of the intermediate duct 80 (where “open cross-sectionalarea of the exhaust duct 70” refers to the cross-sectional area of theexhaust duct 70 not occupied by the intermediate duct 80). In aparticular embodiment, the ratio of the diameter of the intermediateduct 80 on the diameter of the exhaust duct 70 is from 0.2 to 0.4, forexample around ⅓. Other values are also possible, depending for exampleon the optimisation of the weight and cost of the auxiliary power unit10 as a whole.

In the embodiment shown, the intermediate duct 80 is concentric with theperipheral wall 70′ of the exhaust duct 70; the flow of exhaust gases isthus discharged along a central axis C of the exhaust duct 70.

In a particular embodiment, the larger and cooler cooling air flowsurrounding the exhaust gas flow allows for the peripheral wall 70′ ofthe exhaust duct 70 to be made of materials requiring a lower resistanceto high temperature than materials which would be in direct contact withthe exhaust gas flow, where “resistance to high temperature” refers tothe ability of a material to keep their strength, rigidity anddurability when submitted to high temperatures. This may allow for theuse of less expensive materials for the peripheral wall 70′ of theexhaust duct 70. In a particular embodiment, the temperature of the flowagainst the peripheral wall 70′ of the exhaust duct 70 is lower thanthat against the exhaust duct of a traditional gas turbine engineauxiliary power unit, so that the use of high temperature materials(e.g. nickel or titanium alloy) is not required for the peripheral wall70′. For example, the temperature of the exhaust gases may be 800° F. ormore, potentially up to 1200° F.−1400° F., while the cooling air flowtemperature may be 250° F. or less; surrounding the exhaust gas flowwith the cooling air flow thus significantly reduces the temperature ofthe flow in contact with the peripheral wall 70′. In a particularembodiment, the peripheral wall 70′ of the exhaust duct 70 is made ofany suitable aluminum alloy, any suitable light metal alloy, anysuitable composite material including, but not limited to, carbon fibercomposite materials, or any suitable type of polymer.

In a particular embodiment, the fan 78 can be designed to deliver enoughkinetic energy to act as an ejector pump for the exhaust from theturbines 26, 22 and increase the energy delivered by the turbines 26,22.

In particular embodiment, the exhaust of the turbine section isconfigured so that the flow of exhaust gases expelled from theintermediate duct 80 has a higher velocity than the surrounding coolingair flow circulating in the exhaust duct 70. In a particular embodiment,the difference in velocity is selected to create an entrainment effectin the cooling air flow, so as to help circulation of the cooling airflow through the heat exchangers 18, 66, 68 driven by the fan 78. Thismay allow for the size of the fan 78 to be reduced, as compared to aconfiguration without such an entrainment effect.

In a particular embodiment, the inlet and exhaust of the auxiliary powerunit 10 are located on the aircraft skin such that the inlet rampressure significantly exceeds the static pressure at the exhaust plane;this pressure may be used with a venturi effect to depress the staticpressure at the exhaust plane of the turbines 26, 22 in flight, and/orthe fan 78 may be reversible such that it can act as a turbine andrecover energy in high ram conditions where it is not needed to boostcooling flow.

In a particular embodiment, the auxiliary power unit inlet 14 at theaircraft fuselage is provided with a door to prevent unintendedwind-milling and drag when the auxiliary power unit is not operating.Where high speed performance is required in flight this door can beshaped to act as a ram air scoop.

In a particular embodiment, additional aircraft thrust is gained or thedrag penalty is reduced by taking credit for the waste thermal energytransferred to the cooling. In order to maximise this effect (comparableto the Meredith effect in liquid cooled propulsion engines) the sizingof the outlet 82 of the intermediate duct 80 is optimized and theexhaust vector set to provide the maximum propulsive benefit to theaircraft.

Referring to FIG. 1, in a particular embodiment the auxiliary power unit10 includes mounts 84 on the gearbox 28 and near the inlet 74 of theexhaust duct 70; a single inlet flange and a single exhaust flange areprovided for ease of mounting. The integrated cooling system alsofacilitates installation of the auxiliary power unit 10 in thecompartment 76.

FIGS. 5-8 and 11 show an auxiliary power unit 110 in accordance withanother embodiment, where elements similar to that of the embodiment ofFIGS. 1-3 are identified with the same reference numerals and will notbe further described herein.

In this embodiment, the engine core cooler 166 and the generator/enginelubricant cooler 168 are disposed in parallel with respect to oneanother. As can be best seen in FIG. 7, a cooling air duct 186 extendsradially outwardly around a circumference of the exhaust duct 70. Thecooling air duct 186 has an outlet in fluid communication with theexhaust duct 70 and an inlet disposed radially outwardly of the outletand in fluid communication with the compartment 76 through the coolers166, 168. The engine core cooler 166 and the generator/engine lubricantcooler 168 each extend around a respective portion of a circumference ofthe cooling air duct 186. The fan 78 is located in the exhaust duct 70,thus downstream of the coolers 166, 168. As can be seen from FIG. 6, thetwo coolers 166, 168 together extend around only part of thecircumference of the exhaust duct 70, with the intermediate duct 80 andexcess air duct 29 extending adjacent the exhaust duct 70 in thecircumferential portion free of the coolers 166, 168. The coolers 166,168 can be mounted directly to the auxiliary power unit 110 as shown, orcould alternately be installed on the aircraft and linked to theauxiliary power unit 110 with tubing (e.g. flexible tubing).

Referring back to FIG. 7, it can be seen that the air passages 166 b,168 b of the coolers 166, 168 extend along a radial direction R of theauxiliary power unit 110. Alternately, other orientations for thecoolers 166, 168 are possible.

Still referring to FIG. 7, variable pitch blades or variable inlet guidevanes 188 can be provided in the cooling air duct 186 and its junctionwith the exhaust duct 70, immediately upstream of the fan 78, so as tobe able to modulate the airflow through the coolers 166, 168 and/orcontrol fan power absorption at lower heat load conditions.

As can be best seen in FIGS. 5-6, the intercooler 118 is not in fluidcommunication with the exhaust duct 70, and is instead configured as anair to liquid cooler; the intercooler 118 includes fluid passagesreceiving the coolant from the engine core 12′ through one or moreconduits 118′ (for example at about 200° F.) and circulating the coolantin heat exchange relationship with the compressed air from thecompressor 120 (for example at 450° F.) before the coolant is circulatedto the engine core cooler 166 through one or more conduits 118″. Theintercooler 118 is thus located upstream of the engine core cooler 166and downstream of the engine core 12′ in the coolant circulation path.

As can be best seen in FIGS. 6 and 11, in this embodiment twocompressors are provided: a supercharger compressor 120 to providecompressed air to the engine core 12′, and a bleed compressor 121 toprovide bleed air for the aircraft. The two compressors 120, 121 areconnected to the same shaft 124, which also receives the turbines 26, 22of the turbine section. The compressor inlets can be connected to acommon plenum 119 (FIG. 11) or to a respective plenum 119 a, 119 b(FIGS. 5-6, dotted lines in FIG. 11), with the plenum(s) 119, 119 a, 119b being connected to the main inlet 14. In a particular embodiment, sucha configuration allows for accommodating different functionalrequirements for the supercharging flow (to the engine core 12′) and theaircraft flow (to the bleed duct 27).

FIGS. 9-10 show an auxiliary power unit 210 similar to that of FIGS.5-8, where elements similar to that of the embodiment of FIGS. 1-3and/or to that of the embodiment of FIGS. 5-8 are identified with thesame reference numerals and will not be further described herein. Thecompartment 76 is shown as defined by the tail cone 290 of the aircraft,with the exhaust duct outlet 72 located at the tip of the tail cone 290.The tail cone 290 defines the main inlet 14 to the compartment 76, towhich the compressor inlets are connected. The auxiliary power unit ofFIGS. 1-3 and/or of FIGS. 5-8 may be similarly installed.

The engine core cooler 266 and the generator/engine lubricant cooler 268have a rectangular configuration and are circumferentially and axiallyoffset from one another about the exhaust duct 70; each is connected tothe exhaust duct 70 through a respective cooling air duct 286 (FIG. 10)extending radially outwardly from the exhaust duct 70. One or both ofthe coolers 266, 268 can have air conduits angled with respect to theradial direction of the auxiliary power unit 210.

FIG. 12 shows an alternate configuration for the cooling inlet and heatexchangers 318, 366, 368, which may be used in any of the auxiliarypower units 10, 110, 210 described above. A bifurcated inlet systemincludes two separate cooling air ducts 386 a, 386 b, which in aparticular embodiment may allow minimizing the length of the cooling airducts 386 a, 386 b and/or of the coolant/lubricant conduits connected tothe coolers 366, 368 and/or of the compressed air conduits connectingthe intercooler 318 to the compressor 320 and to the engine core 12′.The cooling air duct 386 a closest to the engine core inlet manifold 392is dedicated to the intercooling function and accordingly receives theintercooler 318, which is this embodiment is air cooled. The othercooling air duct 386 b receives one or both of the engine core cooler366 and the generator/engine lubricant cooler 368. The position of theheat exchangers within the cooling air ducts 386 a, 386 b (e.g. how theheat exchangers are grouped in each cooling air duct) may vary, forexample depending on the relative demand for cooling air. The pressurelosses in each cooling air duct 386 a, 386 b of the bifurcated systemare balanced to avoid distorting the inlet flow of the fan 78, which islocated in the exhaust conduit 70 downstream of the heat exchangers 318,366, 368. In a particular embodiment, the generator/engine lubricantcooler 368 is positioned in the same cooling air duct 386 a as theintercooler 318, with the engine core cooler 366 located in the secondcooling air duct 386 b. In another particular embodiment, a whole or apart of the engine core cooler 366 is positioned in the same cooling airduct 386 a as the intercooler 318, with the generator/engine lubricantcooler 368 located in the second cooling air duct 386 b.

FIG. 13 shows another alternate configuration for the cooling inlet andheat exchangers 418, 466, 468, which may be used in any of the auxiliarypower units 10, 110, 210 described above. A bifurcated cooling air duct486 extends non perpendicularly and at a non-zero angle with respect tothe exhaust conduit 70, with an outlet of the cooling air duct 486 beingin fluid communication with the exhaust conduit 70 upstream of the fan78. The heat exchangers are received in the cooling air duct, with theengine core cooler 466 and generator/engine lubricant cooler 468 beinglocated upstream of the intercooler 418. In a particular embodiment, theheat exchangers 418, 466, 468 are placed as close to the engine core 12′as possible, and weight, volume and losses associated with piping thecycle air as well as the lubricant and liquid coolant is minimized.

In a particular embodiment, having the heat exchangers 166, 168, 266,268, 318, 366, 368, 418, 466, 468 located upstream of the fan 78 allowsfor the heat exchangers to be smaller, since the air circulatedtherethrough is cooler. However, the fan 78 downstream of the heatexchangers is exposed to warmer air than a fan upstream of the heatexchangers, and accordingly the power requirement for the fan 78downstream of the heat exchangers may be greater.

FIG. 14 shows an alternate configuration for the two compressors, whichmay be used in replacement of the compressor(s) of any of the auxiliarypower units 10, 110, 210 described above. The supercharger compressor520 providing the compressed air to the engine core 12′ and the bleedcompressor 521 providing the compressed air to the aircraft are arrangedon both sides of a single rotor 594, which in a particular embodiment ismanufactured by forging. The rotor 594 may be received on a shaft 524driven by the turbine section. Tip seals 596 (e.g. labyrinth or fin typeair seals) with a low pressure “sink” (exhaust) 596 below either of theimpeller delivery pressures (e.g. to ambient) are arranged at theimpeller tips to prevent interference between the two compressors 520,521 which might result in premature stall or surge, when the two sidesare operating at different pressures.

FIGS. 15-16 show an alternate configuration for the compressors andturbines, which may be used in replacement of the compressor(s) andturbines of any of the auxiliary power units 10, 110, 210 describedabove. The supercharger compressor 620 is mounted on a separateturbocharger shaft 698 with the second stage (e.g. pressure) turbine622, and where the first stage turbine 626 drives the bleed compressor621 through a turbine shaft 624 and is compounded with the engine core12′ through the gearbox 28. In a particular embodiment, such aconfiguration allows for the turbocharger 620 to find its own matchpoint and possibly eliminate the need for variables on one of thecompressors 620, 621. Variable nozzle geometry (e.g. variable areaturbine vanes 699, see FIG. 16) could be introduced on the second stageturbine 622 to improve controllability of the degree of supercharge. Ina particular embodiment, such a configuration allows for the speed ofthe second stage turbine 622 to be selected independent of therequirements for the first stage turbine 626. As can be seen in FIG. 16,in a particular embodiment the turbocharger shaft 698 is concentric withthe shaft 624 of the first stage turbine 622, and a common inlet plenum619 is provided for both compressors 620, 621. It is understood thatalthough the second stage turbine 622 is shown as a radial turbine, itcould alternately be an axial turbine.

Size effects, material capability and cost considerations generallylimit the efficiency of typical present gas turbine engine auxiliarypower units. In a particular embodiment, the auxiliary power unit 10,110, 210 including some measure of constant volume combustion aided byvariable supercharging to preserve high altitude performance providesfor an increase in efficiency with minimal complexity or need forsophisticated materials requirements and/or improved specific cost ascompared to a traditional gas turbine engine auxiliary power unit.

Like typical auxiliary power unit installations, the auxiliary powerunit 10, 110, 210 can be used to provide both medium pressure air foraircraft use and constant speed shaft power to drive a generator, forexample at synchronous speed for 400 Hz. The auxiliary power unit 10,110, 210 may be operated for air alone, electrical power alone or somecombination of both types of load at the same time. Normally combinedload occurs in ground or low altitude operation. In flight, at altitudesup to the aircraft ceiling, the auxiliary power unit is typicallyrequired to be operable for electrical power only, as an additionalelectrical power source after the main engine(s). In a particularembodiment, the present auxiliary power unit 10, 110, 210 includesvariable supercharging to sustain the required power output in the lessdense air at high altitude.

In a particular embodiment, the auxiliary power unit 10, 110, 210 isconfigured with simple inlet and exhaust connections (including main,load and cooling gas paths) to facilitate quick removal and replacementcomparable to the traditional gas turbine engine auxiliary power units.

It is understood that the engine assemblies shown as auxiliary powerunits 10, 110, 210 may alternately be configured as other types ofengine assemblies, including, but not limited to, turboshaft engineassemblies where the engine core 12′ is configured as or drivinglyengaged to an output shaft, and turboprop engine assemblies where theengine core 12′ is drivingly engaged to a propeller.

The above description is meant to be exemplary only, and one skilled inthe art will recognize that changes may be made to the embodimentsdescribed without departing from the scope of the invention disclosed.Each rotor shown may be a centrifugal or axial device, and may bereplaced by two or more rotors having radial, axial or mixed flowblades. Still other modifications which fall within the scope of thepresent invention will be apparent to those skilled in the art, in lightof a review of this disclosure, and such modifications are intended tofall within the appended claims.

1. A method of cooling a generator and an internal combustion engine ofan auxiliary power unit for an aircraft, the method comprising:circulating a first liquid coolant through the internal combustionengine; circulating a second liquid coolant through the generator; anddriving a cooling air flow in heat exchange relationship with the firstand second liquid coolants using a fan driven by the internal combustionengine.
 2. The method as defined in claim 1, wherein the first coolantis distinct from any fuel and lubricating system of the auxiliary powerunit.
 3. The method as defined in claim 1, wherein the internalcombustion engine is a Wankel rotary engine including a rotor havingthree apex portions mounted for eccentric revolutions within an internalcavity defined in a housing, the internal cavity having an epitrochoidshape with two lobes.
 4. The method as defined in claim 1, wherein thefan is driven by the internal combustion engine through a mechanicalengagement between the fan and the internal combustion engine.